1. Field of the Invention
This invention relates to gas turbine engine combustors having improved arrangements for the provision of air to the interior of the combustion chambers, especially improved arrangements for the provision of that air required for the cooling of the combustion products in the downstream zones within the combustion chamber but not solely these.
2. Discussion of Prior Art
The commercial drive for improved gas turbine engine performance, especially in aircraft engines, renders desirable an increase in turbine inlet temperatures for this would convey an increase in cycle efficiency. The scope for increasing engine efficiency by increase in turbine inlet temperature would seem to dwarf that which could be secured by improvements in aerodynamic design in the compressor and turbine sections. However in current generation engines, particularly advanced military aircraft engines, turbine inlet temperatures are already such as to impose a severe restriction on the useful life of turbine section components despite the use of the best of current generation materials and cooling arrangements within the turbine section. Moreover, at the same time as there is a drive for improved engine performance (as measured by specific fuel consumption) there is a parallel drive for improved component life both in current engines as well as those at the development stage. It is obvious that these two goals will be mutually incompatible unless there is some considerable improvement in engine materials or engine design.
One way in which significant improvements might be achieved without revolutionary change to engine designs or step advances in materials, is by providing a more uniform temperature within the turbine entry flow through improvement in the combustor. Certainly, it has been recognised for some time by those in the art that current generation engines (especially those of the annular combustor design) convey to the turbine a flow of gases having significant variations of temperature from point to point within the turbine entry. A consequence of this is that either the temperature capability of turbine section components, or their endurance, is not utilized to the full but wasted by local hotspots or the like which do not contribute in any way to cycle efficiency. Even when the mean turbine entry temperature is not excessive, stationary hot spots can damage individual turbine guide vanes and a poor radial temperature profile (i.e. across the annulus) can cause uneven degradation of aerodynamic components from root to tip. Considerable progress has already been made in the area of combustion chamber design having regard to those features of earlier designs which produced traceable and detrimental results. Despite these improvements made to date it would be typical for a current generation aircraft engine of the annular combustor type to exhibit an overall temperature distribution factor (OTDF) of say 25%. OTDF is a measure of the highest point temperature less the mean of point temperatures. A figure of 25% obviously means there is still room for improvement in this regard. It is likely however that a new approach to this aspect of performance will be required if any significant reduction of the 25% figure is to be secured.
It has been noticed in the art that in addition to those irregularities in temperature within the turbine entry flow of an annular combustor engine which are traceable to particular known origins within the combustor and can be avoided--such as problems caused by disruption of boundary layers--there is a significant degree of variation which has not been ascribed to any known origin. The term `randomness` has been coined to describe these variations. It is known for example that a particular engine might produce a hot spot (or spots) within the turbine entry field which is consistent in location from engine run to engine run and yet another engine made to the same construction might have its own peculiar hot spots different in location or intensity to the first. It has been suggested that one of the main sources of these asymmetries within an engine's turbine entry field is irregularities within the primary zone of the combustion chamber. However, even in experiments with carefully controlled primary zone exit conditions the effect has persisted and this has led to the suspicion that the asymmetries are created within the dilution zone of the combustion chamber.
Before giving further consideration to the source of the problem addressed in this specification, some discussion of the arrangements found in a conventional gas turbine engine combustor is warranted in order to elucidate the background to the problem and to clarify the terminology used herein. In a typical present day gas turbine combustor there is a combustion liner which may be considered as comprising two or three distinct regions each with its own typical configuration. At the forward end of the combustion chamber (i.e. that end adjacent the compressor outlet) there is a region known as the primary zone in which the primary combustion takes place. The primary zone has arrangements for supplying atomised or vaporised fuel and arrangements for supplying air such that a stabilized recirculatory flow is established for the purposes of maintaining the continuous ignition of new reactants on a localised or general level. At the rearward end of the combustion casing there is a region known as the dilution zone in which air is introduced to the interior entirely for the purpose of cooling and regulating the distribution of the hot gases resulting from combustion to a level tolerable by turbine section components. Usually there is also a region intermediate the primary zone and the dilution zone which is called the intermediate or secondary zone in which air is introduced to the interior for the purpose of completing the combustion process for avoidance of smoke and other emissions and for the avoidance of dissociation loss. The boundary between these three zones is more or less distinct according to combustor design and the intermediate zone may not exist in all combustors as a distinct recognisable zone. The combustion liner sits within an air casing and a portion of the compressor delivery air is funnelled into the space between the combustion liner and the air casing from whence it is fed to the combustion chamber in the various zones.
In the dilution zone this compressor delivery air is introduced to the combustion chamber through relatively large holes in the combustion liner with a view to achieving sufficient cooling jet penetration into the crossflow from the forward zones to secure a turbine entry. The source of the air in these cooling jets is of course the compressor delivery air which is flowing along the outside of the combustion liner and within the air casing.
The characteristics of a single jet, such as a dilution jet, issuing transversely into a crossflow, (typically at 60.degree. say), such as the hot gases from the primary and intermediate zones are well established. An analytical and descriptive text covering this topic may be found at page 117 et seq of Gas Turbine Combustion written by A. H. Lefebvre (published by the McGraw Hill Book Company ISBN 0-07-037029-X) and also in a paper presented by the inventors named in this application, at the 23rd AIAA/SAE/ASME/ASEE Joint Propulsion Conference of Jun. 29-Jul. 2, 1987. This paper is available in reprint form from the American Institute of Aeronautics and Astronautics under reference AIAA-87-1827. Both these above works will be mentioned further in this specification.
The single jet which issues from its source has momentum which projects it into the crossflow causing an obstruction to that crossflow and consequent downstream deflection of the jet from its initial trajectory. There is intensive mixing between jet and crossflow which creates a turbulent shear layer around the periphery of the jet. Gas within this shear layer has less momentum in the direction of the jet than that within the core and consequently it suffers more downstream deflection than the core flow at the sides of the jet where it is free to adopt a different trajectory. This leads to a jet section, downstream of the inlet, which is kidney shaped with the lobes on the inlet side of the jet. Within this overall kidney section there is vortical flow in each lobe with the core gas being swept downstream from the forward and lateral edges and recirculating through the middle of the core.
The situation in a real dilution zone where there are multiple jets and interaction between individual jets is not so clearly established. Generally the flow phenomenon will be similar to that of the single jet. However, the jets present a significant obstruction to the crossflow gases and this blockage effect leads to the creation of a sympathetic pattern of double vortex flow within the crossflow gases in the wake of each jet. There are other complications as well as this one. It has been known for some time that in certain circumstances of adverse dilution zone geometry it is possible for vortex flow to occur within the dilution jet at issue. In the inventors above-referenced paper it was demonstrated that this vortex flow can occur on a gross scale with respect to jet size and with varying degrees of intensity. This is a different phenomenon to that causing vortical flows within the jet and its wake after it issues for it is not a consequence of interaction between jet and crossflow being present within the jet as it leaves the hole. This vortex flow within the jet at issue can however effect that subsequent interaction and various simplistic design rules exist for the avoidance of this phenomenon, which are based on the geometry of the dilution zone. A. H. Lefebvre at page 114 of the above referenced book refers to two aspects of this in the following text:
"If the pitch of the dilution holes is greater than the annulus height, a vortex can form in the flow entering the hole; this changes the penetration and mixing characteristics of the dilution air jet. The strength of the vortex depends on the ratio of annulus area, as measured in the plane of the holes, to the hole area".
In Lefebvre's terminology the annulus is the space between the combustion liner and the surrounding air casing and annulus height is the radial distance between one and the other. Lefebvre reports also (at the same page) that: "Vortex formation, which can occur on both tubular and annular liners, may be eliminated or subdued by fitting a longitudinal splitter plate across (longitudinally with respect to the liner) each dilution hole. The plate . . . is effective when used in conjunction with spectacle plate or dam" (behind the dilution hole within the annulus). The above given quotations correspond to established thinking in the art. The established design aim has been to avoid dilution jet vortex flow by providing adequate annulus height if this is possible or failing this to suppress vortex development by use of splitter plates. However spurious irregularities in the temperature traverse of turbine entry flow have persisted despite this approach. Lefebvre at page 7 of his book gives one disadvantage of the annular combustor design as "Difficult to maintain stable outlet temperature traverse" albeit he seems to ascribe this to difficulty in maintaining a steady velocity profile in the inlet gases (compressor outlet flow).
A recent investigation of the inventors (published in their aforementioned paper) has shed new light on the flow and mixing phenomena within a realistic multiple jet dilution zone model. The model had simplified geometry representative of a typical present day annular combustion zone but with only single sided dilution. The model was manufactured to an extremely tight tolerance so as to remove physical irregularities as a source of flow disturbance. Furthermore, great care was taken to ensure that both the approach flow in the dilution zone and the feedflow to the dilution holes was uniform and consistent to the limits of measurement. This reported investigation demonstrated that there was vortex flow with the dilution jets at issue and confirmed that this vortex flow did significantly influence subsequent events within the dilution zone. The presence of this vortex flow is consistent with Lefebvre's hole pitch to annulus height criterion for the relevant dimensions of the model were 69.85 mm and 35.8 mm respectively. It was however unexpected because the model was representative of real engine designs and presumably not predicted by the design rules on which these real engines had been based. Moreover it was found that the dilution jet internal vortex flow varied in configuration, strength, and rotational sense from hole to hole around the dilution zone annulus. It was noticed that there was aerodynamic twisting of the jets from certain dilution holes at stations downstream of the dilution hole annulus which gave rise to circumferential asymmetry in the temperature distribution, and that the double vortex structure within the jet core (caused by interaction with the crossflow) usually consisted of vortices of unequal strength. It was concluded that (amongst other things) there was evidence to suggest that the direction and location of the vortices formed in the holes influences the rate of mixing between jet and crossflow fluid.
The nature of the in-hole vortex flow revealed by this published investigation suggests that it is not solely dictated by the overall geometry of the dilution zone and the presence of consistent hole to hole differences in a precision made model would seem to indicate that the basic dilution zone arrangements are not sufficiently aerodynamically stable to withstand minuscule irregularities still less those likely to be encountered in a production engine once subjected to thermal cycling in service.